The present invention generally relates to repair of structures fabricated from composite materials and, more particularly, to a process method for repairing aircraft structures fabricated from high temperature materials, such as graphite or fiberglass bismaleimide (BMI) materials.
High performance aircraft, such as military fighters, are typically constructed using composite materials with a higher temperature capability than the composite materials used in other aircraft, such as ordinary commercial aircraft and private aircraft. These higher temperature composite materials, including graphite or fiberglass BMI materials, are being applied to wing, fuselage, and empennage structural components to provide the strength, stiffness, and temperature capability required to fulfill the aircraft""s mission. Unfortunately, damage to these structural components is not easily repaired unless a bolted repair can be applied. Bolted repairs, however, will not properly restore the form, fit, and function of many structural components. If a bolted repair is inappropriate, a bonded-on patch repair is required.
The ideal repair of a structural component is one that fully restores its structural integrity to what it was before it was damaged. The best way to achieve full restoration is to use the same materials and processes for repair as those used for initial fabrication. To achieve the required structural and temperature capability in BMI composite materials, the repair material must be processed at a temperature of 375xc2x0 F. for four hours under high pressure, followed by an additional 6-hour cure at 440xc2x0 F. The high pressure required is achieved by placing the repair material in a pressurized vessel called an autoclave. Maintenance facilities, however, typically do not have autoclave equipment, relying instead upon vacuum bag technology for curing and bonding of composite repair materials.
Standard vacuum bag repair processes will not produce quality BMI repair patches as the standard process traps air and volatiles from the resin and allows the loss of too much resin into vacuum bagged consumable materials. The resulting repair is resin starved, contains an excessive amount of porosity and voids, and does not have sufficient load-carrying capability. When processing BMI materials, the resin""s viscosity becomes extremely low in the 230-280xc2x0 F. temperature range. The low resin viscosity poses a problem for vacuum bagged processes as excessive resin can bleed very easily into the bagging materials unless a dam system is constructed. Since fiber reinforced BMI impregnated material is fabricated with a very low resin content, any reduction of resin can result in degraded structural properties of the material. Conversely, an air path from the BMI material to the vacuum source is necessary to remove gaseous volatiles and air entrapped within the repair material during the cure process. Prevention of excessive resin bleeding, for example, by incorporating a dam in the vacuum bag, counteracts provision of an adequate air path for proper curing. As vacuum pressure will not suppress the formation of voids as positive pressure can, the BMI repair material resulting from the standard vacuum bag repair process contains excessive porosity and void content, and in some cases, delaminates to the point of falling apart.
In addition, many repairs must be performed while the structural component is installed on the aircraft because the structural component either isn""t removable from the airframe or the aircraft operations do not permit removal of the structural component. Consequently, the process of exposing the structural component to the high processing temperature, approximately 375-440xc2x0 F., can be a safety hazard when performed on the aircraft because, for example, the high processing temperature may exceed the temperature allowed in areas on or near the aircraft containing fuel vapors. Damage to other materials near the repair site may also occur due to the high processing temperature. The general guideline due to these hazards is to limit on-aircraft repair processing temperatures to 350xc2x0 F. or less.
U.S. Pat. No. 5,958,166, issued Sep. 28, 1999, entitled xe2x80x9cMethod For Repairing High Temperature Composite Structuresxe2x80x9d discloses repair processes using standard, albeit high temperature, vacuum bagging processes for polyimides such as PMR-15 and AFR700B. U.S. Pat. No. 5,618,606, issued Apr. 8, 1997, entitled xe2x80x9cProcess For Bonding Staged Composites With A Cobonded Staged Adhesive And Articlexe2x80x9d discloses a method to stage epoxy based materials using a non-autoclave vacuum and heating tool to draw off volatiles and partially cure the material for use later. U.S. Pat. No. 5,236,646, issued Aug. 17, 1993, entitled xe2x80x9cProcess For Preparing Thermoplastic Compositesxe2x80x9d discloses a process using a dual vacuum chamber apparatus for consolidating fiber-reinforced thermoplastic prepregs into laminates or composites. The prior art, however, addresses neither the problems related to BMI material consolidation described above, such as prevention of excessive resin bleeding and provision of adequate air path for proper curing, nor issues related to on-aircraft repair described above, such as safety and heat damage to other materials near the repair site.
As can be seen, there is a need for a repair process method specifically to repair composite aircraft structures containing BMI resin. There is also a need for a repair process method that can be performed with flight line maintenance personnel and equipment, to contour and cure a BMI repair patch. Moreover, there is a need for a repair process method that can be used to safely and effectively make on-aircraft repairs containing BMI resin.
The present invention provides a repair process method specifically to repair composite aircraft structures containing BMI resin. The present invention also provides a repair process method that applies a patch consolidation and transfer process, which can be performed with flight line maintenance personnel and equipment, to contour and cure a BMI repair patch. Moreover, the present invention provides a repair process method that can be used to safely and effectively make on-aircraft repairs by bonding a BMI repair patch onto the aircraft structure using conventional bonded repair procedures.
In one aspect of the present invention, a method for repairing a damaged area, referred to as the repair area, of a composite structure comprises steps of:
making alignment markings on the repair area;
fabricating a pair of alignment templates;
preparing the repair area for a hot bonded, vacuum bagged repair;
assembling a repair patch;
consolidating the repair patch;
heating the repair patch;
transferring and aligning the repair patch to the repair area;
vacuum bagging, heating, and cooling the repair patch at the repair area;
heating and cooling the repair patch in an oven; and
bonding the repair patch to the repair area.
In another aspect of the present invention, a method for repairing a damaged area, referred to as the repair area, of a composite structure comprises steps of:
making alignment markings on the repair area;
fabricating a pair of alignment templates;
preparing the repair area for a hot bonded, vacuum bagged repair;
assembling a repair patch;
consolidating the repair patch;
heating the repair patch;
transferring and aligning the repair patch to the repair area;
vacuum bagging, heating, and cooling the repair patch at the repair area;
heating and cooling the repair patch in an oven; and
bonding the repair patch to the repair area.
The step of preparing the repair area for a hot bonded, vacuum bagged repair further includes:
installing thermocouples in the repair area;
removing a flash tape in a bond area of the repair area and leaving the flash tape outside of the bond area; and
covering the repair area with a solid separator film;
The step of assembling a repair patch further includes:
cutting to size plies of fiber-reinforced BMI repair material;
fabricating a ply lay-up template;
taping a first of the pair of alignment templates on the ply lay-up template;
laying up the plies on the alignment template; and
affixing a second of the pair of alignment templates to the repair material;
The step of transferring and aligning the repair patch to the repair area further includes:
transferring the repair patch to the repair area;
aligning a remaining one of the pair of alignment templates with the alignment markings on the repair area;
rotating the repair patch to contact the solid separator film covering the repair area; and
removing the remaining one of the pair of alignment templates from the repair patch;
In still another aspect of the present invention, a method for repairing a damaged area, referred to as the repair area, of a composite structure comprises steps of:
making alignment markings on the repair area;
fabricating a pair of alignment templates;
preparing the repair area for a hot bonded, vacuum bagged repair;
assembling a repair patch;
consolidating the repair patch;
heating the repair patch;
transferring and aligning the repair patch to the repair area;
vacuum bagging, heating, and cooling the repair patch at the repair area;
heating and cooling the repair patch in an oven; and
bonding the repair patch to the repair area.
The step of fabricating a pair of alignment templates further includes centering a piece of transparent bagging film over the repair area; and tracing the alignment markings onto the transparent bagging film.
The step of preparing the repair area for a hot bonded, vacuum bagged repair further includes:
installing thermocouples in the repair area;
removing a flash tape in a bond area of the repair area and leaving the flash tape outside of the bond area; and
covering the repair area with a solid separator film;
The step of assembling a repair patch further includes:
cutting to size plies of fiber-reinforced BMI repair material;
fabricating a ply lay-up template;
taping a first of the pair of alignment templates on the ply lay-up template;
laying up the plies on the alignment template; and
affixing a second of the pair of alignment templates in an alignment template assembly to the repair material.
The step of consolidating the repair patch further includes:
covering a center area of a plate with insulation;
covering the insulation with a heating blanket;
placing a thin copper sheet over the heating blanket;
covering the copper sheet with a non-porous separator film;
transferring the repair patch onto the non-porous separator film, largest ply down;
removing one of the pair of alignment templates from the repair patch;
placing thermocouples around the repair patch;
placing a porous separator film onto the repair patch;
placing a pricked non-porous separator film onto the porous separator film;
placing a first fiberglass cloth over the pricked non-porous separator film;
vacuum bagging the insulation, the heating blanket, the copper sheet, the non-porous separator film, the repair patch with alignment template assembly, the thermocouples, the porous separator film, the pricked non-porous separator film, and the first fiberglass cloth using a lower vacuum bag;
applying a first vacuum to the lower vacuum bag;
placing a second fiberglass cloth onto the lower vacuum bag;
placing a rigid box onto the second fiberglass cloth so that the rigid box is substantially centered over the repair patch;
placing a third fiberglass cloth over the rigid box;
vacuum bagging the rigid box using an upper vacuum bag; and
applying a second vacuum to the upper vacuum bag, so that the second vacuum is at a level ranging between approximately zero inches to one inches of mercury less than the first vacuum;
The step of heating the repair patch further includes:
heating the repair patch at a rate of 3xc2x0 F. per minute to approximately 250xc2x0 F.;
holding the repair patch at approximately 250xc2x0 F. for approximately 15 minutes;
releasing the second vacuum from the upper vacuum bag, opening the upper vacuum bag, and removing the rigid box; the third fiberglass cloth, and the second fiberglass cloth;
holding the repair patch at approximately 250xc2x0 F. for approximately 15 minutes; and
removing the repair patch from the lower vacuum bag;
The step of transferring and aligning the repair patch to the repair area further includes:
transferring the repair patch to the repair area;
aligning the alignment template assembly with the alignment markings on the repair area;
rotating the repair patch to contact the solid separator film covering the repair area; and
removing the alignment template assembly from the repair patch;
The step of vacuum bagging, heating, and cooling the repair patch at the repair area further includes:
placing a precut perforated separator film over the repair patch;
placing a fourth fiberglass cloth over the precut perforated separator film;
placing a second solid separator film over the fourth fiberglass cloth;
placing a heat blanket over the second solid separator film;
placing a fifth fiberglass cloth over the heat blanket;
vacuum bagging the repair patch;
heating the repair patch at a rate of 3xc2x0 F. per minute to approximately 350xc2x0 F.;
holding the repair patch at approximately 350xc2x0 F. for approximately 3 hours; and
cooling the repair patch at a rate of 3xc2x0 F. per minute to approximately ambient temperature;
The step of heating and cooling the repair patch in an oven further includes:
marking orientation on the repair patch;
removing the repair patch from the repair area;
placing the repair patch in an oven;
heating the repair patch at a rate of 3xc2x0 F. per minute to approximately 375xc2x0 F.;
holding the repair patch at approximately 375xc2x0 F. for approximately 3 hours;
heating the repair patch at a rate of 3xc2x0 F. per minute to approximately 440xc2x0 F.;
holding the repair patch at approximately 440xc2x0 F. for approximately 6 hours; and
cooling the repair patch at a rate of 3xc2x0 F. per minute to approximately ambient temperature.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.